U.S. patent application number 15/502311 was filed with the patent office on 2017-08-17 for gas turbine.
The applicant listed for this patent is MITSUBISHI HITACHI POWER SYSTEMS, LTD.. Invention is credited to Shinya HASHIMOTO, Keita TAKAMURA, Masanori YURI.
Application Number | 20170234135 15/502311 |
Document ID | / |
Family ID | 55399376 |
Filed Date | 2017-08-17 |
United States Patent
Application |
20170234135 |
Kind Code |
A1 |
TAKAMURA; Keita ; et
al. |
August 17, 2017 |
GAS TURBINE
Abstract
An axial-direction passage, a forced vortex passage, a first
blade array passage, and a second blade array passage are formed in
a rotor shaft of a turbine. Cooling air from an air extraction port
of a compressor flows through the axial-direction passage which
extends in an axial direction. The forced vortex passage is
connected to the axial-direction passage and extends outwards in
the radial direction relative to an axial line from a connecting
portion between the forced vortex passage and the axial-direction
passage. The first blade array passage is connected to an end
portion on the outer side in the radial direction of the forced
vortex passage and guides cooling air to a first blade row among a
plurality of blade rows. The second blade array passage is
connected to an end portion of the forced vortex passage and guides
cooling air to a second blade row.
Inventors: |
TAKAMURA; Keita;
(Yokohama-shi, JP) ; YURI; Masanori;
(Yokohama-shi, JP) ; HASHIMOTO; Shinya;
(Yokohama-shi, JP) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
MITSUBISHI HITACHI POWER SYSTEMS, LTD. |
Kanagawa |
|
JP |
|
|
Family ID: |
55399376 |
Appl. No.: |
15/502311 |
Filed: |
July 28, 2015 |
PCT Filed: |
July 28, 2015 |
PCT NO: |
PCT/JP2015/071418 |
371 Date: |
February 7, 2017 |
Current U.S.
Class: |
415/116 |
Current CPC
Class: |
F01D 5/08 20130101; F01D
5/084 20130101; F02C 7/18 20130101; Y02T 50/676 20130101; Y02T
50/673 20130101; F05D 2260/213 20130101; Y02T 50/60 20130101; F01D
9/065 20130101; F01D 5/085 20130101; F02C 7/185 20130101; F01D 5/18
20130101; F02C 6/08 20130101 |
International
Class: |
F01D 5/18 20060101
F01D005/18; F02C 7/18 20060101 F02C007/18; F01D 5/08 20060101
F01D005/08 |
Foreign Application Data
Date |
Code |
Application Number |
Aug 29, 2014 |
JP |
2014-175185 |
Claims
1. A gas turbine, comprising: a compressor that compresses air; a
combustor that combusts fuel in the air compressed by the
compressor to generate combustion gas; and a turbine that is driven
by the combustion gas, the compressor including an air extraction
port that extracts air from an intermediate stage of the compressor
as cooling air, the turbine including a rotor shaft, and a
plurality of blade rows that are arranged in an axial direction of
the rotor shaft and each include a plurality of blades attached to
an outer circumference of the rotor shaft, an axial-direction
passage, a forced vortex passage, a first blade array passage, and
a second blade array passage being formed in the rotor shaft, the
axial-direction passage being connected to the air extraction port
and extending in the axial direction, the forced vortex passage
being connected to the axial-direction passage and extending
outwards in a radial direction of the rotor shaft from a connecting
portion between the forced vortex passage and the axial-direction
passage, the first blade array passage being connected to an end
portion on an outer side in the radial direction of the forced
vortex passage and configured to guide the cooling air to a first
blade row among the plurality of blade rows, the second blade array
passage being connected to an end portion on the outer side in the
radial direction of the forced vortex passage and configured to
guide the cooling air to a second blade row among the plurality of
blade rows.
2. The gas turbine according to claim 1, wherein: the forced vortex
passage is formed on a downstream side of a third blade row among
the plurality of blade rows, the third blade row being disposed on
a furthest downstream side of a flow of the combustion gas in the
axial direction.
3. The gas turbine according to claim 1, further comprising: an air
extraction pipe that connects the air extraction port and the
axial-direction passage; an air extraction branch pipe connected to
the air extraction pipe; and a pre-swirl nozzle that is connected
to the air extraction branch pipe and imparts a speed component in
a rotational direction of the rotor shaft to the cooling air that
flowed through the air extraction branch pipe, a third blade array
passage that guides the cooling air that passed through the
pre-swirl nozzle to the third blade row, among the plurality of
blade rows, disposed on the furthest downstream side of the flow of
the combustion gas in the axial direction being formed in the rotor
shaft.
4. The gas turbine according to claim 1, wherein: the compressor
includes a second air extraction port that extracts air from an
intermediate stage further on an upstream side of the flow of the
air inside the compressor than a first air extraction port, serving
as the air extraction port, as cooling air; and a second
axial-direction passage, a second forced vortex passage, and a
third blade array passage are formed in the rotor shaft, the second
axial-direction passage being connected to the second air
extraction port and extending in the axial direction on the outer
side in the radial direction with respect to a first
axial-direction passage serving as the axial-direction passage, the
second forced vortex passage being connected to the second
axial-direction passage, and extending outwards in the radial
direction from the connecting portion between the second forced
vortex passage and the second axial-direction passage on the
downstream side of the third blade row, among the plurality of
blade rows, disposed on the furthest downstream side of the flow of
the combustion gas in the axial direction, the third blade array
passage being connected to an end portion on the outer side in the
radial direction of the second forced vortex passage, and
configured to guide the cooling air to the third blade row.
5. The gas turbine according to claim 2, further comprising: an air
extraction pipe that connects the air extraction port and the
axial-direction passage; an air extraction branch pipe connected to
the air extraction pipe; and a pre-swirl nozzle that is connected
to the air extraction branch pipe and imparts a speed component in
a rotational direction of the rotor shaft to the cooling air that
flowed through the air extraction branch pipe, a third blade array
passage that guides the cooling air that passed through the
pre-swirl nozzle to the third blade row, among the plurality of
blade rows, disposed on the furthest downstream side of the flow of
the combustion gas in the axial direction being formed in the rotor
shaft.
6. The gas turbine according to claim 2, wherein: the compressor
includes a second air extraction port that extracts air from an
intermediate stage further on an upstream side of the flow of the
air inside the compressor than a first air extraction port, serving
as the air extraction port, as cooling air; and a second
axial-direction passage, a second forced vortex passage, and a
third blade array passage are formed in the rotor shaft, the second
axial-direction passage being connected to the second air
extraction port and extending in the axial direction on the outer
side in the radial direction with respect to a first
axial-direction passage serving as the axial-direction passage, the
second forced vortex passage being connected to the second
axial-direction passage, and extending outwards in the radial
direction from the connecting portion between the second forced
vortex passage and the second axial-direction passage on the
downstream side of the third blade row, among the plurality of
blade rows, disposed on the furthest downstream side of the flow of
the combustion gas in the axial direction, the third blade array
passage being connected to an end portion on the outer side in the
radial direction of the second forced vortex passage, and
configured to guide the cooling air to the third blade row.
Description
TECHNICAL FIELD
[0001] The present invention relates to a gas turbine.
[0002] This application claims priority based on Japanese Patent
Application No, 2014-17,518,5 filed in Japan on Aug. 29, 2014, of
which the contents are incorporated herein by reference.
BACKGROUND ART
[0003] A gas turbine includes a compressor that compresses
atmospheric air to generate compressed air, a combustor that
combusts fuel in this compressed air to generate combustion gas,
and a turbine that is driven by the combustion gas. The turbine
includes a turbine rotor that rotates around an axial line, and a
turbine casing that covers the turbine rotor. The turbine rotor
includes a rotor shall that extends in an axial direction in which
the axial line extends, centered around the axial line, and a
plurality of blade rows fixed to this rotor shaft. The plurality of
blade rows are arranged in the axial direction. Each of the blade
rows includes a plurality of blades arranged in a circumferential
direction with the axial line serving as a reference.
[0004] The blades of the turbine come into contact with
high-temperature combustion gas, and are therefore often cooled by
some method. For example, in Patent Document 1 below, air extracted
from the compressor is utilized as cooling air for the blades. In
the rotor shaft described in Patent Document 1, a rotor bore tube,
a first forced vortexing passage, a second forced vortexing
passage, a first passage, and a second passage are formed. The
rotor bore tube is a cavity, elongated in the axial direction, into
which cooling air from the compressor flows. The first forced
vortexing passage extends outwards in the radial direction from the
rotor bore tube at a position in the axial direction between a
first stage blade row and a second stage blade row. The second
forced vortexing passage extends outwards in the radial direction
from the rotor bore tube at a position in the axial direction
between the second stage blade row and a third stage blade row. The
first passage guides the cooling air that passed through the first
forced vortexing passage to the first stage blade row. The second
passage guides the cooling air that passed through the second
forced vortexing passage to the second stage blade row.
CITATION LIST
Patent Document
[0005] Patent Document 1: Japanese Unexamined Patent Application
Publication No. 2009-275705A
SUMMARY OF THE INVENTION
Technical Problems
[0006] According to the technology described in Patent Document 1 a
passage extending outwards in the radial direction that supplies
the cooling air from the compressor is formed for each of the
plurality of blade rows. Thus, according to the technology
described in Patent Document 1, the rotor shall is elongated,
resulting in problems such as a degradation in vibration
characteristics of the turbine rotor and a reduction in aerodynamic
performance of the turbine.
[0007] Hence, an object of the present invention is to provide a
gas turbine capable of cooling a blade using cooling air from a
compressor while, at the same time, suppressing a degradation in
vibration characteristics of a rotor.
Solution to Problems
[0008] A gas turbine according to a first aspect of the present
invention for solving the above-described problems includes a
compressor that compresses air, a combustor that combusts fuel in
the air compressed by the compressor to generate combustion gas,
and a turbine that is driven by the combustion gas. The compressor
includes an air extraction port that extracts air from an
intermediate stage of the compressor as cooling air. The turbine
includes a rotor shall, and a plurality of blade rows that are
arranged in an axial direction of the rotor shaft and each include
a plurality of blades attached to an outer circumference of the
rotor shaft. An axial-direction passage, a forced vortex passage, a
first blade array passage, and a second blade array passage are
formed in the rotor shaft. The axial-direction passage is connected
to the air extraction port and extends in the axial direction. The
forced vortex passage is connected to the axial-direction passage
and extends from a connecting portion between the forced vortex
passage and the axial-direction passage outwards in a radial
direction with respect to the rotor shaft. The first blade array
passage is connected to an end portion on an outer side in the
radial direction of the forced vortex passage and guides the
cooling air to a first blade row among a plurality of blade rows.
The second blade array passage is connected to an end portion on
the outer side in the radial direction of the forced vortex passage
arid guides the cooling air to a second blade row among the
plurality of blade rows.
[0009] According to the gas turbine, the number of forced vortex
passages that are formed in the rotor shaft and extend in the
radial direction can be decreased. Thus, according to the gas
turbine, it is possible to suppress elongation of the rotor shaft
associated with formation of the forced vortex passages, and thus
suppress deterioration of vibration characteristics of the rotor.
Further, according to the gas turbine, it is possible to suppress
extension of a distance between the plurality of stages associated
with formation of the forced vortex passages, and thus suppress a
reduction in aerodynamic performance of the turbine.
[0010] Further, in a gas turbine according to a second aspect of
the present invention for solving the above-described problems, the
forced vortex passage may be formed on a downstream side of a third
blade row among the plurality of blade rows, the third blade row
being disposed on a furthest downstream side of a flow of the
combustion gas in the axial direction. In this case, the first
blade row and the second blade row are provided on an upstream side
of the third blade row furthest downstream among the plurality of
blade rows.
[0011] According to the gas turbine, it is possible to suppress
extension of the distance between all the stages. Further,
according to the gas turbine, it is possible to access the end
portion on the outer side in the radial direction of the forced
vortex passage without disassembling the rotor shaft. Thus,
according to the gas turbine, even if a foreign substance such as
debris accumulates at the end portion on the outer side in the
radial direction of the forced vortex passage, it is possible to
easily collect the foreign substance.
[0012] Further, the gas turbine according to the first or second
aspect may further include an air extraction pipe that connects the
air extraction port and the axial-direction passage, an air
extraction branch pipe connected to the air extraction pipe, and a
pre-swirl nozzle that is connected to the air extraction branch
pipe and imparts a speed component in a rotational direction of the
rotor shaft to the cooling air that flowed through the air
extraction branch pipe. A third blade array passage that guides the
cooling air that passed through the pre-swirl nozzle to the third
blade row, among the plurality of blade rows, disposed on the
furthest downstream side of the flow of the combustion gas in the
axial direction may be formed in the rotor shaft.
[0013] According to the gas turbine, the cooling air that cools the
first blade row and the second blade row is introduced into the
forced vortex passage, causing an increase in pressure.
Nevertheless, according to the gas turbine, the cooling air that
cools the third blade row is passed through the pre-swirl nozzle
without being introduced into the forced vortex passage, causing
the cooling air to swirl and thus achieving a reduction in a
rotation resistance of the turbine rotor. Thus, according to the
gas turbine, the air extracted from a single air extraction port is
used as cooling air, making it possible to simplify the cooling air
system and, at the same time, decrease the rotation resistance of
the turbine rotor and increase efficiency of the gas turbine.
[0014] Further, in the gas turbine according to the first or second
aspect, a second air extraction port may be formed in the
compressor, and a second axial-direction passage, a second forced
vortex passage, and the third blade array passage may be formed in
the rotor shaft. The second air extraction port extracts air from
an intermediate stage further on an upstream side of the flow of
the air inside the compressor than a first air extraction port,
which is the air extraction port, as cooling air. The second
axial-direction passage is a passage into which the cooling air
from the second air extraction port flows, and extends in the axial
direction on the outer side in the radial direction with respect to
a first axial-direction passage, which is the axial-direction
passage. The second forced vortex passage is connected to the
second axial-direction passage, and extends outwards in the radial
direction from the connecting portion between the second forced
vortex passage and the second axial-direction passage on the
downstream side of the third blade row, among the plurality of
blade rows, disposed on the furthest downstream side of the flow of
the combustion gas in the axial direction. The third blade array
passage is connected to an end portion on the outer side in the
radial direction of the second forced vortex passage, and guides
the cooling air to the third blade row.
[0015] Thus, cooling air having a pressure lower than a pressure of
the cooling air supplied to the first blade row and the second
blade row can be supplied to the third blade row disposed furthest
downstream among a plurality of vane rows. Moreover, according to
the gas turbine, the air extracted from the compressor is
introduced into the rotating rotor shaft, and a forced vortex of
the cooling air is produced in this rotor shaft, thereby increasing
the pressure of the cooling air. This air is then supplied to the
third blade row. As a result, according to the gas turbine, it is
possible to extract air having a pressure lower than that of the
air from the first air extraction port from the second air
extraction port as cooling air, and utilize this air as the cooling
air for the third blade row. Thus, according to the gas turbine, a
compression ratio of the cooling air for the third blade row can be
decreased by the compressor, making it possible to suppress a
driving force of the compressor and increase the efficiency of the
gas turbine.
Advantageous Effects of Invention
[0016] According to one aspect of the present invention, it is
possible to cool a blade using cooling air from a compressor and,
at the same time, suppress elongation of a rotor shaft and thus
suppress a degradation in vibration characteristics of a rotor.
BRIEF DESCRIPTION OF THE DRAWINGS
[0017] FIG. 1 is an overall cutaway side view of a main portion of
a gas turbine of a first embodiment according to the present
invention.
[0018] FIG. 2 is a cross-sectional view of the main portion of the
gas turbine of the first embodiment according to the present
invention.
[0019] FIG. 3 is an explanatory view illustrating a flow of cooling
air and a state quantity thereof in the first embodiment according
to the present invention.
[0020] FIG. 4A and 4B illustrate a rotating body in which a forced
vortex passage is formed, FIG. 4A being a front view of the
rotating body and FIG. 4B being a cross-sectional view thereof
taken along the line B-B of FIG. 4A.
[0021] FIG. 5 is a cross-sectional view of a main portion of a gas
turbine of a second embodiment according to the present
invention.
[0022] FlG. 6 is a perspective view of a main portion of a
pre-swirl nozzle of the second embodiment according to the present
invention.
[0023] FIG. 7 is an explanatory view illustrating a flow of cooling
air and a state quantity thereof in the second embodiment according
to the present invention.
[0024] FIG. 8 is a cross-sectional view of a main portion of a gas
turbine of a third embodiment according to the present
invention.
[0025] FIG. 9 is a cross-sectional view taken along the line VIII
of FIG. 8.
[0026] FIG. 10 is an explanatory view illustrating a flow of
cooling air and a state quantity thereof in the third embodiment
according to the present invention.
[0027] FIG. 11 is a cross-sectional view of a main portion of a gas
turbine of a modification of the third embodiment according to the
present invention.
DESCRIPTION OF EMBODIMENTS
[0028] The following describes in detail various embodiments of a
gas turbine according to the present invention, with reference to
the drawings.
First Embodiment
[0029] The following describes a first embodiment of the gas
turbine according to the present invention with reference to FIGS.
1 to 4.
[0030] A gas turbine according to the present embodiment includes a
compressor 10 that compresses air, a combustor 20 that combusts a
fuel in the air compressed by the compressor 10 to generate
combustion gas, a turbine 30 that is driven by the combustion gas,
and an air extraction line 80 that feeds the air extracted from the
compressor 10 to the turbine 30 as cooling air, as illustrated in
FIG. 1.
[0031] The compressor 10 includes a compressor rotor 11 that
rotates around an axial line Ar, a compressor casing 15 that covers
the compressor rotor 11, and a plurality of vane rows 14, as
illustrated in FIGS. 1 and 2. Note that, in the following, the
direction in which the axial line Ar extends is referred to as an
axial direction Da, and one side and the other side of this axial
direction Da are referred to as an upstream side and a downstream
side, respectively. This upstream side is an upstream side of a
flow of the air inside the compressor 10 as well as an upstream
side of a flow of the combustion gas inside the turbine 30. Thus,
this downstream side is a downstream side of the flow of the air
inside the compressor 10 as well as a downstream side of the flow
of the combustion gas inside the turbine 30. Further, a
circumferential direction around this axial line Ar is simply
referred to as a circumferential direction Dc, and a direction
orthogonal to the axial line Ar is referred to as a radial
direction Dr. The compressor rotor 11 includes a rotor shaft 12
that is centered around the axial line Ar and extends in the axial
direction Da, and a plurality of blade rows 13 attached to this
rotor shaft 12. The plurality of blade rows 13 are arranged in the
axial direction Da. Each of the blade rows 13 includes a plurality
of blades arranged in the circumferential direction Dc. The vane
rows 14 are respectively disposed on the downstream side of the
plurality of blade rows 13. Each of the vane rows 14 is provided on
an inner side of the compressor casing 15. Each of the vane rows 14
includes a plurality of vanes arranged in the circumferential
direction Dc. An annular space between an outer peripheral side in
the radial direction of the rotor shaft 12 and an inner peripheral
side in the radial direction of the compressor casing 15 in a
region where the vane rows 14 and the blade rows 13 are disposed in
the axial direction Da forms an air compression flow channel 19
through which air flows and, at the same time, is compressed. That
is, this compressor 10 is an axial flow multistage compressor. In
the compressor casing 15, a medium-pressure air extraction port 16
is formed in a position corresponding, to an intermediate
stage.
[0032] The turbine 30 includes a turbine rotor 31 that rotates
around the axial line Ar, a turbine casing 35 that covers the
turbine rotor 31, and a plurality of vane rows 34. The combustor 20
is fixed to a section on the upstream side of this turbine casing
35. The turbine rotor 31 includes a rotor shaft 32 that is centered
around the axial line Ar and extends in the axial direction Da, and
a plurality of blade rows 33 attached to this rotor shaft 32. The
plurality of blade rows 33 are arranged in the axial direction Da.
Each of the blade rows 33 includes a plurality of blades arranged
in the circumferential direction Dc. The vane rows 34 are
respectively disposed on the upstream side of the plurality of
blade rows 33. Each of the vane rows 34 is provided on an inner
side of the turbine easing 35. Each of the vane rows 34 includes a
plurality of vanes arranged in the circumferential direction Dc. A
combustion gas flow channel 39 through which combustion gas G from
the combustor 20 flows is formed in an annular space between an
outer peripheral side of the rotor shaft 32 and an inner peripheral
side of the turbine casing 35 in a region where the vane rows 34
and the blade rows 33 are disposed in the axial direction Da.
[0033] The turbine 30 of the present embodiment has four stages.
Thus, the turbine 30 of the present embodiment includes a first
stage vane row 34a, a second stage vane row 34b, a third stage vane
row 34c, and a fourth stage vane row 34d as the vane rows 34.
Further, the turbine 30 of the present embodiment includes a first
stage blade row 33a, a second stage blade row 33b, a third stage
blade row 33c, and a fourth stage blade row 33d as the blade rows
33.
[0034] The combustor 20 is fixed to a section on the upstream side
of the turbine casing 35. This combustor 20 includes a combustion
liner (or transition piece) 21 that feeds the high-temperature,
high-pressure combustion gas G into the combustion gas flow channel
39 of the turbine 30, and a fuel injector 22 that injects fuel
along with air compressed by the compressor 10 into this combustion
liner 21.
[0035] The compressor rotor 11 and the turbine rotor 31 are
positioned on the same axial line Ar and connected with each other
to form a gas turbine rotor 1. Further, the compressor casing 15
and the turbine casing 35 are connected with each other to form a
gas turbine casing 5.
[0036] The air extraction line 80 includes a medium-pressure air
extraction pipe 81, a cooler 86 provided to this medium-pressure
air extraction pipe 81, an air adjustment valve 87 provided to the
medium-pressure air extraction pipe 81, and an air regulator 84
provided to the medium-pressure air extraction pipe 81. The
medium-pressure air extraction pipe 81 includes a medium-pressure
air extraction main pipe 82 connected to the medium-pressure air
extraction port 16 of the compressor 10, and a vane medium-pressure
air extraction pipe 83 as well as a blade medium-pressure air
extraction pipe 85 connected to the medium-pressure air extraction
main pipe 82.
[0037] The vane medium-pressure air extraction pipe 83 is connected
to a position corresponding to the intermediate stage of the
turbine 30 on the turbine casing 35. The air regulator 84, such as
an orifice, for regulating a pressure and a flow rate of the air
that passes through this vane medium-pressure air extraction pipe
83 is provided to the vane medium-pressure air extraction pipe 83.
Note that this air regulator 84 may be an adjustment valve. The
blade medium-pressure air extraction pipe 85 is connected to the
rotor shaft 32. The cooler 86 and the air adjustment valve 87
described above are provided to this blade medium-pressure air
extraction pipe 85. The cooler 86 cools the air that passes through
this blade medium-pressure air extraction pipe 85. The air
adjustment valve 87 adjusts the flow rate of the air that passes
through this blade medium-pressure air extraction pipe 85. Note
that the cooler 86 may be provided to the medium-pressure air
extraction main pipe 82.
[0038] The rotor shaft 32 of the turbine rotor 31 includes a
plurality of rotor discs 42, as illustrated in FIG. 2. The
plurality of rotor discs 42 are arranged in the axial direction Da
and connected to each other by a spindle bolt 41 that is inserted
through the rotor disks 42 in the axial direction Da. The turbine
30 of the present embodiment includes a first disc 42a, a second
disc 42b, a third disc 42c, and a fourth disc 42d as the rotor
discs 42. One blade row 33 is attached to each of the plurality of
rotor discs 42. That is, the first stage blade row 33a is attached
to the first disc 42a, the second stage blade row 33b is attached
to the second disc 42b, the third stage blade row 33c is attached
to the third disc 42c, and the fourth stage blade row 33d is
attached to the fourth disc 42d.
[0039] The rotor shaft 32 includes a small diameter portion 43
supported by a hearing 70, and a large diameter portion 44 that has
a larger outer diameter than an outer diameter of the small
diameter portion 43. The plurality of blade rows 33 are attached to
an outer circumference of this large diameter portion 44. The small
diameter portion 43 is provided on the downstream side of the large
diameter portion 44. An outer peripheral side of the hearing 70 is
covered by a bearing cover 71. An upstream side seal member 77a
that seals an area between the hearing cover 71 and the small
diameter portion 43 of the rotor shaft 32 is provided on the
upstream side of the bearing 70, and a downstream side seal member
72b that seals an area between the bearing cover 71 arid the small
diameter portion 43 of the rotor shaft 32 is provided on the
downstream side of the bearing 70, on an inner peripheral side of
the bearing cover 71.
[0040] An inner diffuser 77 having a cylindrical shape and centered
around the axial line Ar is disposed on the outer peripheral side
of the small diameter portion 43, and an outer diffuser 78 having a
cylindrical shape and centered around the axial line Ar is disposed
on an outer peripheral side of this inner diffuser 77, on the
downstream side of the large diameter portion 44 of the rotor
shaft. 31 The inner diffuser 77 and the outer diffuser 78 are both
directly or indirectly fixed to the turbine casing 35. An annular
space between the outer peripheral side of the inner diffuser 77
and the inner peripheral side of the outer diffuser 78 forms a
combustion gas exhaust now path 79 through which the combustion gas
flowing out from the combustion gas flow channel 39 flows.
[0041] An axial-direction passage 45, a forced vortex passage 46, a
second stage blade array passage 47 (first blade array passage), a
third stage blade array passage 48 (second blade array passage),
and a fourth stage blade array passage 49 (third blade array
passage) are formed in the rotor shaft 32. The axial-direction
passage 45 is a passage into which cooling air from the
medium-pressure air extraction port 16 flows, and extends in the
axial direction Da. The forced vortex passage 46 is connected to
the axial-direction passage 45 and extends outwards in the radial
direction. The second stage blade array passage 47 guides the
cooling air that passed through the forced vortex passage 46 to the
second stage blade row 33b (first blade row). The third stage blade
array passage 48 guides the cooling air that passed through the
forced vortex passage 46 to the third stage blade row 33c (second
blade row). The fourth stage blade array passage 49 guides the
cooling air that passed through the forced vortex passage 46 to the
fourth stage blade row 33d (third blade row). The axial-direction
passage 45 opens on a downstream end of the small diameter portion
43, and extends in the axial direction Da to a position on a
downstream portion of the large diameter portion 44. This
axial-direction passage 45 is a circular column-shaped passage
centered around the axial line Ar. The second stage blade array
passage 47, the third stage blade array passage 48, the fourth
stage blade array passage 49, and the forced vortex passage 46 are
each formed in the large diameter portion 44 of the rotor shaft 32.
The forced vortex passage 46 is formed in a position on the
downstream side of the fourth stage blade row 33d furthest
downstream in the large diameter portion 44. The second stage blade
array passage 47, the third stage blade array passage 48, and the
fourth stage blade array passage 49 are each connected to an end
portion on the outer side in the radial direction of the forced
vortex passage 46.
[0042] Note that, in the specification and claims of the present
application, "A and B are connected" refers to a state in which A
and B are configured so that air flows from A to B or from B to
A.
[0043] Further, in the specification and claims of the present
application, "forced vortex passage" refers to a flow channel that
serves as a passage of a fluid provided to a rotating body and
feeds the fluid outwards or inwards in the radial direction while
swirling the fluid at the same peripheral speed as that of the
rotating body. Such a flow channel is generally a hole H that is
linearly provided in a radial direction of a rotating body R, such
as a rotor disc, and interconnects different positions in the
radial direction, such as illustrated in FIGS. 4A and 48. However,
the forced vortex passage is not limited thereto, and may be a hole
having a curved shape, and may be formed using a cylindrical member
that is attached to a disc and extends in the radial direction or a
blade member that protrudes from a disc in the axial direction as
in a radial compressor.
[0044] The passages in the large diameter portion 44 of the rotor
shaft 32 will now be described in more detail.
[0045] A first cavity 52a is formed between the first disc 42a and
the second disc 42b. A second cavity 52b is formed between the
second disc 42b and the third disc 42c. A third cavity 52c is
formed between the third disc 42c and the fourth disc 42d. A fourth
cavity 52d is formed in a section on the downstream side in the
fourth disc 42d. The first cavity 52a, the second cavity 52b, the
third cavity 52c, and the fourth cavity 52d are each an annular
space with the axial line Ar serving as the center. The fourth
cavity 52d is formed in a position where the forced vortex passage
46 is formed in the axial direction Da, and is connected to the end
portion on the outer side in the radial direction of this forced
vortex passage 46. A fourth disc passage 53d and a fourth stage
communicating passage 57 are formed in the fourth disc 42d. The
fourth disc passage 53d extends in the axial direction Da and
communicates with the fourth cavity 52d and the third cavity 52c.
The fourth stage communicating passage 57 communicates with the
fourth cavity 52d and an attachment position of the fourth stage
blade row 33d. A third disc passage 53c and a third stage
communicating passage 56 are formed in the third disc 42c. The
third disc passage 53c extends in the axial direction Da and
communicates with the second cavity 52b and the third cavity 52c.
The third stage communicating passage 56 communicates with the
second cavity 52b and an attachment position of the third stage
blade row 33c. A second stage communicating passage 55 that
communicates with the second 52b and an attachment position of the
second stage blade row 33b is formed in the second disc 42b.
[0046] The fourth stage blade array passage 49 is formed by the
fourth cavity 52d and the fourth stage communicating passage 57.
The third stage blade array passage 48 is formed by the fourth
cavity 52d, the fourth disc passage 53d, the third cavity 52c, the
third disc passage 53c, the second cavity 52b, and the third stage
communicating passage 56. The second stage blade array passage 47
is formed by the fourth cavity 52d, the fourth disc passage 53d,
the third cavity 52c, the third disc passage 53c, the second cavity
52b, and the second stage communicating passage 55.
[0047] Thus, the fourth cavity 52d forms a passage common to the
fourth stage blade array passage 49, the third stage blade array
passage 48, and the second stage blade array passage 47. Further,
the fourth disc passage 53d, the third cavity 52c, the third disc
passage 53c,and the second cavity 52b form a passage common to the
third stage blade array passage 48 and the second stage blade array
passage 47.
[0048] An axial end flange 74 opposing the rotor shaft 32 at a
distance in the axial direction Da is disposed on a downstream end
of the rotor shaft 32. This axial end flange 74 is fixed to the
bearing cover 71. An end portion of the blade medium-pressure air
extraction pipe 85 is fixed to this axial end flange 74. A
through-hole for communicating an interior of the blade
medium-pressure air extraction pipe 85 and the axial-direction
passage 45 formed in the rotor shaft 32 is formed in this axial end
flange 74.
[0049] Next, the operation of the gas turbine described above will
be described.
[0050] The compressor 10 generates compressed air by sucking in and
compressing ambient air. The compressed air generated by the
compressor 10 is partially discharged into the combustion liner 21
via the fuel injector 22 of the combustor 20. Further, fuel from
the fuel injector 22 is it into the combustion liner 21. This fuel
combusts in the compressed air inside the combustion liner 21. As a
result of this combustion, the combustion gas G is generated, and
this combustion gas G flows from the combustion liner 21 into the
combustion gas flow channel 39 of the turbine 30. This combustion
gas G passes through the combustion gas flow channel 39, thereby
rotating the turbine rotor 31.
[0051] The blades of the turbine 30 disposed inside the combustion
gas flow channel 39 are exposed to the high-temperature combustion
gas. Thus, in the present embodiment, the air extracted from the
compressor 10 is supplied as cooling air to the blades that
constitute the second stage blade row 33b, the third stage blade
row 33c, and the fourth stage blade row 33d, and cools the blades.
Furthermore, in the present embodiment, the cooling air is also
supplied to the vanes that constitute the third stage vane row 34c,
and cools the vanes.
[0052] FIG. 3 illustrates an exemplary temperature and pressure
balance in the gas turbine. The pressure of the air in the
medium-pressure air extraction port 16 of the compressor 10 is 10
ata. Further, the pressure between the second stage vane row 34b
and the second stage blade row 33b in the combustion gas flow
channel 39 is 8 ata, the pressure between the third stage vane row
34c and the third stage blade row 33c in the combustion gas flow
channel 39 is 6 ata, and the pressure between the fourth stage vane
row 34d and the fourth stage blade row 33d in the combustion gas
flow channel 39 is 2 ata.
[0053] The cooling air extracted from the medium-pressure air
extraction port 16 of the compressor 10 flows through the
medium-pressure air extraction main pipe 82 of the air extraction
line 80. A portion of the cooling air then flows into the vane
medium-pressure air extraction pipe 83 while a remaining portion
flows into the blade medium-pressure air extraction pipe 85. The
cooling air introduced into the vane medium-pressure air extraction
pipe 83 reaches a pressure of 7 ata in a process of passing through
the air regulator 84, is supplied to the plurality of vanes that
constitute the third stage vane row 34c, and cools the plurality of
vanes.
[0054] The cooling air introduced into the blade medium-pressure
air extraction pipe 85 is cooled in a process of passing through
the cooler 86, adjusted in flow rate by the air adjustment valve
87, and then introduced into the axial-direction passage 45 of the
rotor shaft 32. Owing to the pressure loss in the process of
passing through the cooler 86 and the air adjustment valve 87 and
the like, the cooling air immediately prior to introduction into
the axial-direction passage 45 reduces in pressure to about 8 ata.
Further, the cooling air immediately prior to introduction into the
axial-direction passage 45 is cooled by the cooler 86, reducing in
temperature.
[0055] The cooling air introduced into the axial-direction passage
45 flows through the forced vortex passage 46 that extends outwards
in the radial direction from this axial-direction passage 45, and
into the fourth cavity 52d. In the process of passing through the
forced vortex passage 46 that extends outwards in the radial
direction, the cooling air is subjected to a centrifugal force from
the rotor shaft 32 that rotates around the axial line Ar,
increasing in pressure. As a result, the pressure of the cooling
air that reached the fourth cavity 52d is 9 ata. Note that the
cooling air in the fourth cavity 52d rises in temperature due to
increased pressure.
[0056] A portion of the cooling air in the fourth cavity 52d flows
through the fourth stage communicating passage 57 that partially
forms the fourth stage blade array passage 49, and into the cooling
air passages in the plurality of blades that constitute the fourth
stage blade row 33d. In the process of passing through the fourth
stage communicating passage 57, this cooling air is regulated in
pressure and flow rate. As a result, the cooling air immediately
prior to flowing from the fourth stage communicating passage 57
into the blades of the fourth stage blade row 33d reaches a
pressure of 3 ata. This cooling air passes through the cooling air
passages in the plurality of blades that constitute the fourth
stage blade row 33d, cools the blades, and then is discharged into
the combustion gas flow channel 39.
[0057] The remaining portion of the cooling air in the fourth
cavity 52d flows through the fourth disc passage 53d, the third
cavity 52c, and the third disc passage 53c, and into the second
cavity 52b. The pressure of the cooling air in the second cavity
52b is approximately the same as the pressure of the cooling air in
the fourth cavity 52d at 9 ata.
[0058] A portion of the cooling air introduced into the second
cavity 52b flows through the third stage communicating passage 56
that partially forms the third stage blade array passage 48,
and,into the cooling air passages in the plurality of blades that
constitute the third stage blade row 33c. In the process of passing
through the third stage communicating passage 56, this cooling air
is regulated in pressure and flow rate, and reaches a pressure of
approximately 7 ata immediately prior to flowing from the third
stage communicating passage 56 into the blades of the third stage
blade row 33c. This cooling air passes through the cooling air
passages in the plurality of blades that constitute the third stage
blade row 33c, cools the blades, and then is discharged into the
combustion gas flow channel 39.
[0059] The remaining portion of the cooling air introduced into the
second cavity 52b flows through the second stage communicating
passage 55 that partially forms the second stage blade array
passage 47, and into the cooling air passages in the plurality of
blades that constitute the second stage blade row 33b. In the
process of passing through the second stage communicating passage
55, this cooling air is regulated in pressure and flow rate. As a
result, the cooling air immediately prior to flowing from the
second stage communicating passage 55 into the blades of the second
stage blade row 33b reaches a pressure of approximately 9 ata. This
cooling air passes through the cooling air passages in the
plurality of blades that constitute the second stage blade row 33b,
cools the blades, and then is discharged into the combustion gas
flow channel 39.
[0060] The cooling air that flows from the fourth cavity 52d
through the fourth stage communicating passage 57 and into the
blades of the fourth stage blade row 33d, the cooling air that
flows from the second cavity 52b through the third stage
communicating passage 56 and into the blades of the third stage
blade row 33c, and the cooling air that flows from the second
cavity 52b through the second stage communicating passage 55 and
into the blades of the second stage blade row 33b each incur a
pressure loss in the process of passing through the communicating
passages 57, 56, 55, and, increase in pressure upon being subjected
to the centrifugal force from the rotor shaft 32. The cooling air
that flows from the fourth cavity 52d through the fourth stage
communicating passage 57 and into the blades of the fourth stage
blade row 33d incurs a greater pressure loss when passing through
the communicating passage 57 than the pressure effect from the
centrifugal three in the process of passing through the fourth
stage communicating passage 57 and, as a result, is reduced in
pressure. Further, the cooling air that flows from the second
cavity 52b through the third stage communicating passage 56 and
into the blades of the third stage blade row 33c also incurs a
greater pressure loss when passing through the communicating
passage 56 than the pressure effect from the centrifugal force and,
as a result, is reduced in pressure. On the other hand, the cooling
air that flows from the second cavity 52b through the second stage
communicating passage 55 and into the blade of the second stage
blade row 33b is suppressed from incurring a pressure loss in the
process of passing through the second stage communicating passage
55.
[0061] As described above, in the present embodiment, it is
possible to cool the blades of the turbine 30 by the air extracted
from the compressor 10. Moreover, in the present embodiment, the
air extracted from the compressor 10 is introduced into the
rotating rotor shaft 32 and a forced vortex of the cooling air is
produced in this rotor shaft 32, thereby increasing the pressure of
the cooling air. This cooling air is then supplied to each of the
blade rows 33. As a result, in the present embodiment, it is
possible to extract low-pressure air from the compressor 10, making
it possible to suppress the driving, force of the compressor 10.
Thus, in the present embodiment, it is possible to cool the blades
of the turbine 30 by the air extracted from the compressor 10 and,
at the same time, suppress a reduction in the efficiency of the gas
turbine.
[0062] Further, in the present embodiment, to suppress the pressure
of the air extracted from the compressor 10, the forced vortex
passage 46 that extends in the radial direction is formed its the
rotor shaft 32, and the air extracted from the compressor 10 is it
into the forced vortex passage 46, increased in pressure, and then
distributed to the blades of each of the blade rows 33. Note that
the forced vortex passage 46 may be formed for each of the
plurality of blade rows 33. Nevertheless, in this case, the
plurality of forced vortex passages 46 are formed inside the rotor
shaft 32 in different positions in the axial direction Da and, as a
result, the length of the rotor shaft 32 in the axial direction Da
increases, degrading the vibration characteristics of the turbine
rotor 31. Furthermore, the distance between stages of the turbine
30 also increases, reducing the aerodynamic performance of the
turbine 30.
[0063] On the other hand, in the present embodiment as described
above, once the air is increased in pressure in the forced vortex
passage 46, the air is distributed to each of the plurality of
blade rows 33, thereby suppressing elongation of the rotor shaft 32
in the axial direction Da and thus making it possible to suppress
deterioration of the vibration characteristics of the turbine rotor
31. Furthermore, the forced vortex passage 46 is formed on the
downstream side of the fourth stage blade row 33d furthest
downstream and does not exist in any location between the plurality
of stages, making it possible to suppress elongation of the
distance between the stages of the turbine 30 and thus suppress a
reduction in the aerodynamic performance of the turbine 30 as
well.
[0064] Further, a foreign substance such as debris contained in the
air that flows through the forced vortex passage 46 is expected to
accumulate at the end portion on the outer side in the radial
direction of the forced vortex passage 46. In the present
embodiment, this forced vortex passage 46 is formed in a position
on the downstream side of the fourth stage blade row 33d furthest
downstream, making it possible to access the end portion on the
outer side in the radial direction of this forced vortex passage 46
without disassembling the rotor shaft 32 into the plurality of
rotor discs 42. Thus, in the present embodiment, even if a foreign
substance such as debris accumulates at the end portion on the
outer side in the radial direction of the forced vortex passage 46,
it is possible to easily collect the foreign substance.
Second Embodiment
[0065] The following describes a second embodiment of the gas,
turbine according to the present invention with reference to FIGS.
5 to 7.
[0066] In the gas turbine of the first embodiment, the cooling air
extracted from the medium-pressure air extraction port 16 of the
compressor 10 is separated into air for the second stage blade row
33b, air for the third stage blade row 33c, and air for the fourth
stage blade row 33d in the rotor shaft 32 of the turbine 30.
[0067] In the gas turbine of the present embodiment, the cooling
air extracted from the medium-pressure air extraction port 16 of
the compressor 10 is first separated into air for the second and
third stage blade rows 33b, 33c, and air for the fourth stage blade
row 33d outside the rotor shaft 32 of the turbine 30, as
illustrated in FIGS. 5 and 7. Furthermore, in this gas turbine, the
air for the second and third stage blade rows 33b, 33c is separated
into air for the second stage blade row 33b and air for the third
stage blade row 33c in the rotor shaft 32.
[0068] For this purpose, in the present embodiment, an air
extraction branch pipe 88 that guides a portion of the cooling air
that flowed through the blade medium-pressure air extraction pipe
85 into the fourth stage blade row 33d as air for the fourth stage
blade row 33d is connected to the blade medium-pressure air
extraction pipe 85 connected to the downstream end of the rotor
shaft 32. This air extraction branch pipe 88 is connected to a
downstream end of the large diameter portion 44 of the rotor shaft
32.
[0069] In the rotor shaft 32 of the present embodiment, similarly
to the rotor shaft. 32 of the first embodiment, the axial-direction
passage 45, the forced vortex passage 46, the, second stage blade
array passage 47 (first blade array passage), and the third stage
blade array passage 48 (second blade array passage) are formed. The
axial-direction passage 45 is a passage into which cooling air from
the medium-pressure air extraction port 16 flows via the blade
medium-pressure air extraction pipe 85. The forced vortex passage
46 is connected to the axial-direction passage 45. The second stage
blade array passage 47 guides the cooling air that passed through
the forced vortex passage 46 to the second stage blade row 33b
(first blade row). The third stage blade array passage 48 guides
the cooling air that passed through the forced vortex passage 46 to
the third stage blade row 33c (second blade row).
[0070] The first cavity 52a, the second cavity 52b, the third
cavity 52c, and the fourth cavity 52d are formed in the large
diameter portion 44 of the rotor shaft 32, similarly to the large
diameter portion 44 of the rotor shaft 32 of the first embodiment.
The fourth cavity 52d is connected to the end portion on the outer
side in the radial direction of the forced vortex passage 46. The
fourth cavity 52d and the third cavity 52c are connected by the
fourth disc passage 53d, and the third cavity 52c and the second
cavity 52b are connected by the third disc passage 53c. The second
cavity 52b and the third stage blade row 33c are connected by the
third stage communicating passage 56, and the second cavity 52b and
the second stage blade row 33b are connected by the second stage
communicating passage 55.
[0071] Thus, in the present embodiment as well, the third stage
blade array passage 48 is formed by the fourth cavity 52d, the
fourth disc passage 53d, the third cavity 52c, the third disc
passage 53c, the second cavity 52b, and the third stage
communicating passage 56. Further, the second stage blade array
passage 47 is formed by the fourth cavity 52d, the fourth disc
passage 53d, the third cavity 52c, the third disc passage 53c, the
second cavity 52b, and the second stage communicating passage
55.
[0072] In the large diameter portion 44 of the rotor shall 32 of
the present embodiment, a fifth cavity 52e indented from the
downstream side end of the large diameter portion 44 toward the
upstream side is further formed in a position on the outer side in
the radial direction with respect to the fourth cavity 52d. This
fifth cavity 52e and the fourth stage blade row 33d are connected
by a fourth stage communicating, passage 57a. Thus, in the present
embodiment, a fourth stage blade array passage 49a (third blade
array passage) that guides the cooling air that passed through a
pre-swirl nozzle 67 into the fourth stage blade row 33d is formed
so as to include the fifth cavity 52e and the fourth stage
communicating, passage 57a.
[0073] An axial end plate 61 having a disc shape with the axial
line Ar serving as the center is disposed opposing the downstream
end of the large diameter portion 44 at a distance in the axial
direction Da, on the downstream side of the large diameter portion
44. An end on an inner side in the radial direction of the axial
end plate 61 is fixed to an upstream end of the bearing cover 71. A
large diameter portion end cover 62 is disposed opposing the
downstream end of the large diameter portion 44 at a distance in
the axial direction Da, between this axial end plate 61 and the
large diameter portion 44. An end on the inner side in the radial
direction of this large diameter portion end cover 62 is fixed to
the axial end plate 61. Further, an end on the outer side in the
radial direction of this large diameter portion end cover 62 is
fixed to an upstream end of the inner diffuser 77. A seal
attachment portion 63 opposing an outer peripheral surface on the
downstream side of the large diameter portion 44 at a distance is
formed in a section on the outer side in the radial direction of
the large diameter portion end cover 62. This seal attachment
portion 63 is provided with a seal member 64 that seals an area
between the large diameter portion 44 and the large diameter
portion end cover 62.
[0074] A large diameter portion end cavity 65 that communicates
with the fifth cavity 52e and is an annular space with the axial
line Ar as the center is formed between the large diameter portion
end cover 62 and the large diameter portion 44. Further, an air
receiving space 66 that is an annular space with the axial line Ar
as the center is formed in the region where the fifth cavity 52e is
formed in the radial direction Dr, between the large diameter
portion end cover 62 and the axial end plate 61. This air receiving
space 66 is a space surrounded by the large diameter portion end
cover 62 and the axial end plate 61.
[0075] An end portion of the air extraction branch pipe 88 is fixed
to the axial end plate 61. As a result, a portion of the cooling
air that flowed through the blade medium-pressure air extraction
pipe 85 flows through the air extraction branch pipe 88 and into
the air receiving space 66 formed by the axial end plate 61 and the
large diameter portion end cover 62. In the large diameter portion
end cover 62, the pre-swirl nozzle 67 that imparts a speed
component in the rotational direction of the rotor shaft 32 to the
cooling air introduced into the air receiving space 66 is provided
in a position opposing the fifth cavity 52e of the large diameter
portion 44. In the large diameter portion end cover 62, a seal
member 64b that seals an area between the large diameter portion 44
and the large diameter portion end cover 62 is provided in a
position opposing the large diameter portion 44 in the radial
direction, on the inner side in the radial direction with respect
to the position where the pre-swirl nozzle 67 is provided.
[0076] The pre-swirl nozzle 67 includes a plurality of swirl vanes
68 arranged in the circumferential direction Dc, as illustrated in
FIG. 6. The swirl vane 68 gradually inclines to a rotational
direction Rr side of the rotor shaft 32, from the downstream side
toward the upstream side. This pre-swirl nozzle 67 is a nozzle that
partially converts the pressure of the cooling air introduced into
the air receiving space 66 on the downstream side to kinetic energy
in the rotational direction Rr of the rotor shaft 32, imparting a
speed component in the rotational direction Rr to the cooling
air.
[0077] Next, the flow of the cooling air in the gas turbine
described above will be described.
[0078] The cooling air extracted from the medium-pressure air
extraction port 16 of the compressor 10 and introduced into the
blade medium-pressure air extraction pipe 85 is, similarly to the
first embodiment, cooled in the process of passing through the
cooler 86, adjusted in flow rate by the air adjustment valve 87,
and then partially introduced into the axial-direction passage 45
of the rotor shaft 32. The cooling air immediately prior to
introduction into the axial direction passage 45 reduces in
pressure to about 8 ata as illustrated in FIG. 7, and reduces in
temperature as well.
[0079] The cooling air introduced into the axial-direction passage
45, similarly to the first embodiment, flows through the forced
vortex passage 46 that extends outwards in the radial direction
from this axial-direction passage 45, and into the fourth cavity
52d. In the process of passing through the forced vortex passage 46
that extends outwards in the radial direction, the cooling air is
subjected to a centrifugal force from the rotor shaft 32 that
rotates around the axial line Ar, increasing in pressure. As a
result, the pressure of the cooling air that reached the fourth
cavity 52d is 9 ata. Note that the cooling air in the fourth cavity
52d rises in temperature due to increased pressure.
[0080] The cooling air in the fourth cavity 52d flows through the
fourth disc passage 53d, the third cavity 52c, and the third disc
passage 53c, and into the second cavity 52b. The pressure of the
cooling air in the second cavity 52b is approximately the same as
the pressure of the cooling air in the fourth cavity 52d at 9 ata,
and the temperature of this cooling air is approximately the same
as the temperature of the cooling air in the fourth cavity 52d.
[0081] A portion of the cooling air introduced into the second
cavity 52b, similarly to the first embodiment, flows through the
third stage communicating passage 56 that partially forms the third
stage blade array passage 48, and into the cooling air passages in
the plurality of blades that constitute the third stage blade row
33c. In the process of passing through the third stage
communicating passage 56, this cooling air is regulated in pressure
and flow rate, and reaches a pressure of approximately 7 ata
immediately prior to flowing from the third stage communicating
passage 56 into the blades of the third stage blade row 33c. This
cooling air passes through the cooling air passages in the
plurality of blades that constitute the third stage blade row 33c,
cools the blades, and then is discharged into the combustion gas
flow channel 39.
[0082] The remaining portion of the cooling air introduced into the
second cavity 52b, similarly to the first embodiment, flows through
the second stage communicating passage 55 that partially forms the
second stage blade array passage 47, and into the cooling air
passages in the plurality of blades that constitute the second
stage blade row 33b. In the process of passing through the second
stage communicating passage 55, this cooling air is regulated in
pressure and flow rate. As a result, the cooling air immediately
prior to flowing from the second stage communicating passage 55
into the blades of the second stage blade row 33b reaches a
pressure of approximately 9 ata. This cooling air passes through
the cooling air passages in the plurality of blades that constitute
the second stage blade row 33b, cools the blades, and then is
discharged into the combustion gas flow channel 39.
[0083] A portion of the cooling air that flows through the blade
medium-pressure air extraction pipe 85 flows into the air
extraction branch pipe 88. The pressure of the cooling air
introduced into the air extraction branch pipe 88 is approximately
8 ata.
[0084] The cooling air flows from the air extraction branch pipe 88
into the air receiving space 66, then through the pre-swirl nozzle
67 and into the large diameter portion end cavity 65. In the
process of the cooling air flowing through the pre-swirl nozzle 67,
the pressure of the cooling air is partially converted to kinetic
energy in the rotational direction Rr of the rotor Shaft 32,
imparting a speed component in the rotational direction Rr to the
cooling air. The cooling air upon passing through the pre-swirl
nozzle 67 reduces in pressure to about 5 ata. Further, a
circumferential velocity, with reference to the axial line Ar, of
the cooling air upon passing through the pre-swirl nozzle 67 is
approximately the same as a circumferential velocity in a position
of the fifth cavity 52e in the rotor shaft 32. That is, the cooling
air in the large diameter portion end cavity 65 swirls in the large
diameter portion end cavity 65 having an annular shape at a
circumferential velocity that is approximately the same as that in
the fifth cavity 52e in the rotor shaft 32. Thus, the cooling air
in the large diameter portion end cavity 65 does not cause
resistance that hinders rotation of the rotor shaft 32 when flowing
into the fifth cavity 52e in the rotor shaft 32.
[0085] The pressure of the cooling air introduced into the fifth
cavity 52e is approximately 5 ata.
[0086] The cooling air in the fifth cavity 52e flows through the
fourth stage communicating passage 57a that partially forms the
fourth stage blade array passage 49a, and into the cooling air
passages in the plurality of blades that constitute the fourth
stage blade row 33d. In the process of passing through the fourth
stage communicating passage 57a, this cooling air is, regulated in
pressure and flow rate. As a result, the cooling air immediately
prior to flowing from the fourth stage communicating passage 57a
into the blades of the fourth stage blade row 33d reaches a
pressure of 3 ata. This cooling air passes through the cooling air
passages in the plurality of blades that constitute the fourth
stage blade row 33d, cools the blades, and then is discharged into
the combustion gas flow channel 39.
[0087] Thus, in the present embodiment as well, similarly to the
first embodiment, it is possible to cool the blades of the turbine
30 by the air extracted from the compressor 10. Furthermore, in the
present embodiment as well, it is possible to suppress the driving
force of the compressor 10 and thus suppress a reduction in the
efficiency of the gas turbine.
[0088] Further, in the present embodiment as well, similarly to the
first embodiment, the air is increased in pressure in the forced
vortex passage 46 formed on the downstream side of the blade row 33
furthest downstream, and is then distributed to each of the
plurality of blade rows 33, making it possible to suppress a
deterioration in the vibration characteristics of the turbine rotor
31 as well as a reduction in the aerodynamic performance of the
turbine 30.
[0089] Further, in the first embodiment, the cooling air that cools
the second stage blade row 33b, the third stage blade row 33c, and
the third stage blade row 33c is introduced into the forced vortex
passage 46 in its entirety, and the centrifugal force generated by
the rotation of the turbine rotor 31 is then utilized to increase
the pressure of the cooling air. On the other hand, in the present
embodiment, only the cooling air that cools the second stage blade
row 33b and the third stage blade row 33c is introduced into the
forced vortex passage 46 and increased in pressure, and the cooling
air that cook the fourth stage blade row 33d passes through the re
swirl nozzle 67 without flowing into the forced vortex passage 46,
causing this cooling air to swirl, decreasing a rotation resistance
of the turbine rotor 31. Thus, in the present embodiment, the
rotation resistance of the turbine rotor 31 becomes less than that
in the first embodiment, making it possible to increase the
efficiency of the gas turbine.
[0090] Further, in the present embodiment, the cooling air for the
fourth stage blade row 33d passes through the air extraction branch
pipe 88 provided outside the rotor shaft 32, making it possible to
easily regulate the flow rate, pressure, and temperature of the
cooling air for this fourth stage blade row 33d by providing the
air adjustment valve and the cooler to this air extraction branch
pipe 88.
Third Embodiment
[0091] The following describes a third embodiment of the gas
turbine according to the present invention with reference to FIGS.
8 to 10.
[0092] In the gas turbine of both the first and second embodiments,
the cooling air extracted from the medium-pressure air extraction
port 16 of the compressor 10 is utilized as the cooling air for the
second, third, and fourth stage blade rows 33b, 33c, 33d.
[0093] In the gas turbine of the present embodiment, as illustrated
in FIGS. 8 and 10, the cooling air extracted from the
medium-pressure air extraction port 16 (first air extraction port)
of the compressor 10 is utilized as the cooling air for the
second,and third stage blade rows 33b, 33c. Furthermore, in this
gas turbine, the cooling air extracted from a low-pressure air
extraction port 17 (second air extraction port) of the compressor
10 is utilized as the cooling air for the fourth stage blade row
33d.
[0094] The medium-pressure air extraction port 16 and the
low-pressure air extraction port 17 are formed in positions
corresponding to the intermediate stage of the compressor casing 15
of the present embodiment. The low-pressure air extraction port 17
is formed on the upstream side of the position where the
medium-pressure air extraction port 16 is formed. As a result, the
pressure of the air extracted from the low-pressure air extraction
port 17 is lower than the pressure of the air extracted from the
medium-pressure air extraction port 16, and is 6 atm, for
example.
[0095] An air extraction line 80b of the present embodiment,
similarly to the first embodiment, includes the medium-pressure air
extraction pipe 81 as well as the cooler 86, the air adjustment
valve 87, and the air regulator 84 provided to this medium-pressure
air extraction pipe 81. The medium-pressure air extraction pipe 81,
similarly to the first embodiment, includes the medium-pressure air
extraction main pipe 82 connected to the medium-pressure air
extraction port 16 of the compressor 10, and the vane
medium-pressure air extraction pipe 83 as well as the blade
medium-pressure air extraction pipe 85 connected to the
medium-pressure air extraction main pipe 82. The air regulator 84
described above is provided to the vane medium-pressure air
extraction pipe 83, and the cooler 86 and the air adjustment valve
87 described above are provided to the blade medium pressure air
extraction pipe 85.
[0096] The air extraction line 80b of the present embodiment
further includes a low-pressure air extraction pipe 91 as well as a
cooler 96, an air adjustment valve 97, and an air regulator 94
provided to this low-pressure air extraction pipe 91. The
low-pressure air extraction pipe 91 includes a low-pressure air
extraction main pipe 92 connected to the low-pressure air
extraction port 17 of the compressor 10, and a vane low-pressure
air extraction pipe 93 as well as a blade low-pressure air
extraction pipe 95 connected to the low-pressure air extraction
main pipe 92.
[0097] The vane low-pressure air extraction pipe 93 is connected to
a position corresponding to the intermediate stage of the turbine
30 in the turbine casing 35. More specifically, the vane
low-pressure air extraction pipe 93 is connected on the downstream
side of the position where the vane medium-pressure air extraction
pipe 83 is connected, in the turbine casing 35. The air regulator
94 such as an orifice, for regulating the pressure and the flow
rate of the air that passes through this vane low-pressure air
extraction pipe 93 is provided to the vane low-pressure air
extraction pipe 93. The blade low-pressure air extraction pipe 95
is connected to the rotor shaft 32. The cooler 96 and the air
adjustment valve 97 described above are provided to this blade
low-pressure air extraction pipe 95. The cooler 96 cools the air
that passes through this blade low-pressure air extraction pipe 95.
The air adjustment valve 97 adjusts the flow rate of the air that
passes through this blade low-pressure air extraction pipe 95. An
end portion of this blade low-pressure air extraction pipe 95 is
fixed to the bearing cover 71.
[0098] In the rotor shaft 32 of the present embodiment, similarly
to those in the first and second embodiments, the first
axial-direction passage 45, the first forced vortex passage 46, the
second stage blade array passage 47 (first blade array passage),
and the third stage blade array passage 48 (second blade array
passage) are formed. The first axial-direction passage 45 is a
passage into which cooling air from the medium-pressure air
extraction port 16 flows, and extends in the axial direction Da.
The first forced vortex passage 46 is connected to the first
axial-direction passage 45, and extends outwards in the radial
direction. The second stage blade array passage 47 guides the
cooling air that passed through the first forced vortex passage 46
to the second stage blade row 33b (first blade row). The third
stage blade array passage 48 guides the cooling air that passed
through the first forced vortex passage 46 to the third stage blade
row 33c (second blade row). Furthermore, in the rotor shaft 32 of
the present embodiment, a second axial-direction passage 45b, a
second forced vortex passage 46b, and a fourth stage blade array
passage 49b (third blade array passage) are formed. The second
axial-direction passage 45b is a passage into which cooling air
from the low-pressure air extraction port 17 flows, and extends in
the axial direction Da. The second forced vortex passage 46b is
connected to the second axial-direction passage 45b, and extends
outwards in the radial direction. The fourth stage blade array
passage 49b guides the cooling air that passed through the second
forced vortex passage 46b to the fourth stage blade row 33d (third
blade row).
[0099] The first cavity 52a, the second cavity, the third cavity
52c, and the fourth cavity 52d are formed in the large diameter
portion 44 of the rotor shaft 32, similarly to the large diameter
portion 44 of the rotor shaft 32 of the first and second
embodiments. The fourth cavity 52d is connected to the end portion
on the outer side in the radial direction of the first forced
vortex passage 46. The fourth cavity 52d and the third cavity 52c
are connected by the fourth disc passage 53d, and the third cavity
52c and the second cavity 52b are connected by the third disc
passage 53c. The second cavity 52b and the third stage blade row
33c are connected by the third stage communicating passage 56, and
the second cavity 52b and the second stage blade row 33b are
connected by the second stage communicating passage 55.
[0100] Thus, in the present embodiment as well, the third stage
blade array passage 48 is formed by the fourth cavity 52d, the
fourth disc passage 53d, the third cavity 52c, the third disc
passage 53c, the second cavity 52b, and the third stage
communicating passage 56. Further, the second stage blade array
passage 47 is formed by the fourth cavity 52d, the fourth disc
passage 53d, the third cavity 52c, the third disc passage 53c, the
second cavity 52b, and the second stage communicating passage
55.
[0101] The second axial-direction passage 45b, as illustrated in
FIGS. 8 and 9, is formed in a position on the outer side in the
radial direction with respect to the first axial-direction passage
45 that extends in the axial direction Da with the axial line Ar
serving as the center. A low-pressure air receiving passage 45c is
formed on a downstream end of the second axial-direction passage
45b. This low-pressure air receiving passage 45c extends outwards
in the radial direction from this downstream end, and opens at the
outer peripheral surface of the small diameter portion 43.
[0102] In the bearing cover 71, a through-hole that penetrates the
bearing cover 71 from an outer peripheral side to an inner
peripheral side is formed in substantially the same position as the
opening of the low-pressure air receiving passage 45c in the axial
direction Da. The end portion of the blade low-pressure air
extraction pipe 95 is fixed at the position where this through-hole
is formed. The upstream side and the downstream side of the
through-hole of the bearing cover 71, between the inner peripheral
side of the bearing cover 71 and the outer peripheral side of the
small diameter portion 43, are sealed by seal members 72b, 72c,
respectively.
[0103] The outer side in the radial direction of the large diameter
portion 44, which is the downstream side, is covered by a large
diameter portion end cover 62b. A fifth cavity 52f having an
annular shape with the axial line Ar serving as the center is
formed between the large diameter portion end cover 62b and the
large diameter portion 44.
[0104] The second forced vortex passage 46b that extends outwards
in the radial direction from an upstream end of the second
axial-direction passage 45b and communicates with the fifth cavity
52f is connected to the upstream end of the second axial-direction
passage 45b. This second forced vortex passage 46b is formed in a
position on the downstream side of the first forced vortex passage
46 in the large diameter portion 44. A fourth stage communicating
passage 57b that communicates with the fifth cavity 52f and an
attachment position of the fourth stage blade row 33d is formed in
the large diameter portion 44.
[0105] Thus, the fourth stage blade array passage 49b in the
present embodiment is formed by the fifth cavity 52f and the fourth
stage communicating passage 57b.
[0106] Next, the flow of the cooling air in the gas turbine
described above will be described.
[0107] The cooling air extracted from the medium-pressure air
extraction port 16 of the compressor 10 and introduced into the
blade medium-pressure air extraction pipe 85 is, similarly to the
first and second embodiments, cooled in the process of passing
through the cooler 86, adjusted in flow rate by the air adjustment
valve 87, and then introduced into the first axial-direction
passage 45 of the rotor shaft 32. The cooling air immediately prior
to introduction into the first axial-direction passage 45 reduces
in pressure to about 8 ata illustrated in FIG. 10, and reduces in
temperature as well.
[0108] The cooling air introduced into the first axial-direction
passage 45, similarly to the first and second embodiments, flows
through the first forced vortex passage 46 that extends outwards in
the radial direction from this first axial-direction passage 45,
and into the fourth cavity 52d. In the process of passing through
the forced vortex passage 46 that extends outwards in the radial
direction, the cooling air is subjected to a centrifugal force from
the rotor shaft 32 that rotates around the axial line Ar,
increasing in pressure. As a result, the pressure of the cooling
air that reached the fourth cavity 52d is 9 ata. Note that the
cooling air in the fourth cavity 52d rises in temperature due to
increased pressure.
[0109] The cooling air in the fourth cavity 52d flows through the
fourth disc passage 53d, the third cavity 52c, and the third disc
passage 53c, and into the second cavity 52b. The pressure of the
cooling air in the second cavity 52b is approximately the same as
the pressure of the cooling air in the fourth cavity 52d at 9 ata,
and the temperature of this cooling air is approximately the same
as the temperature of the cooling air in the fourth cavity 52d.
[0110] A portion of the cooling air introduced into the second
cavity 52b, similarly to the first and second embodiments, flows
through the third stage communicating passage 56 that partially
forms the third stage blade array passage 48, and into the cooling
air passages in the plurality of blades that constitute the third
stage blade row 33c. In the process of passing through the third
stage communicating passage 56, this cooling air is regulated in
pressure and flow rate. As a result, the cooling air immediately
prior to flowing from the third stage communicating passage 56 into
the blades of the third stage blade row 33c reaches a pressure of
approximately 7 ata. This cooling air passes through the cooling
air passages in the plurality of blades that constitute the third
stage blade row 33c, cools the blades, and then is discharged into
the combustion gas flow channel 39.
[0111] The remaining portion of the cooling air introduced into,
the second cavity 52b, similarly to the first embodiment, flows
through the second stage communicating passage 55 that partially
forms the second stage blade array passage 47, and into the cooling
air passages in the plurality of blades that constitute the second
stage blade row 33b. In the process of passing through the second
stage communicating passage 55, this cooling air is regulated in
pressure and flow rate. As a result, the cooling air immediately
prior to flowing from the second stage communicating passage 55
into the blades of the second stage blade row 33b reaches a
pressure of approximately 9 ata. This cooling air passes through
the cooling air passages in the plurality of blades that constitute
the second stage blade row 33b, cools the blades, and then is
discharged into the combustion gas flow channel 39.
[0112] The cooling air extracted from the low-pressure air
extraction port 17 of the compressor 10 flows through the
low-pressure air extraction main pipe 92 of the air extraction line
80b. A portion of the cooling air then flows into the vane
low-pressure air extraction pipe 93 while a remaining portion flows
into the blade low-pressure air extraction pipe 95. The cooling air
introduced into the vane low-pressure air extraction pipe 93
reaches a pressure of 4 ata in the process of passing through the
air regulator 94, is supplied to the plurality of vanes that
constitute the fourth stage vane row 34d, and cools the plurality
of vanes.
[0113] The cooling air introduced into the blade low-pressure air
extraction pipe 95 is cooled in the process of passing through the
cooler 96, adjusted in flow rate by the air adjustment valve 97,
and then introduced into the second axial-direction passage 45b of
the rotor shaft 32. Owing to the pressure loss in the process of
passing through the cooler 96 and the air adjustment valve 97 and
the like, the cooling air immediately prior to introduction into
the second axial-direction passage 45b reduces in pressure to about
4 ata. Further, the cooling air immediately prior to introduction
into the second axial-direction passage 45b is cooled by the cooler
86.
[0114] The cooling air introduced into the second axial-direction
passage 45b flows through the second forced vortex passage 46b that
extends outwards in the radial direction from this second
axial-direction passage 45b, and into the fifth cavity 52f. In the
process of passing through the second forced vortex passage 46b
that extends outwards in the radial direction, the cooling air is
subjected to the centrifugal force from the rotor shaft 32 that
rotates around the axial line Ar, increasing in pressure. As a
result, the pressure of the cooling air that reached the fifth
cavity 52f is 5 ata.
[0115] The cooling air in the fifth cavity 52f flows through the
fourth stage communicating passage 57b that partially forms the
fourth stage blade array passage 49b, and into the cooling air
passages in the plurality of blades that constitute the fourth
stage blade row 33d. In the process of passing, through the fourth
stage communicating passage 57b, this cooling air is regulated in
pressure and flow rate. Further, this cooling air increases in an
amount of heat exchange with the heat of the combustion gas as the
cooling air nears the combustion gas flow channel 39, gradually
increasing in temperature. As a result, the cooling air immediately
prior to flowing from the fourth stage communicating passage 57b
into the blades of the fourth stage blade row 33d reaches a
pressure of 3 ata. This cooling air passes through the cooling air
passages in the plurality of blades that constitute the fourth
stage blade row 33d, cools the blades, and then is discharged into
the combustion gas flow channel 39.
[0116] Thus, in the present embodiment as well, similarly to the
first and second embodiments, it is possible to cool the blades of
the turbine 30 by the air extracted from the compressor 10.
Furthermore, in the present embodiment as well, it is possible to
suppress the driving force of the compressor 10 and thus suppress a
reduction in the efficiency of the gas turbine.
[0117] Further, in the present embodiment, the first and second
forced vortex passages 46, 46b are formed on the downstream side of
the blade row 33 furthest downstream, and the air is increased in
pressure in the first forced vortex passage 46 and then distributed
to the second stage blade row 33b and the third stage blade row
33c. As a result, in the present embodiment as well, elongation of
the length of the rotor shaft 32 and the distance between stages is
suppressed, making it possible to suppress a deterioration in the
vibration characteristics of the rotor and a reduction in the
aerodynamic performance of the turbine 30.
[0118] In the present embodiment, cooling air having a pressure
lower than the pressure of the cooling air supplied to the second
stage blade row 33b and the third stage blade row 33c can be
supplied to the fourth stage blade row 33d. Moreover, in the
present embodiment, the air extracted from the compressor 10 is
introduced into the rotating rotor shaft 32 and a forced vortex of
the cooling air is produced in this rotor shaft 32, thereby
increasing the pressure of the cooling air. This air is then
supplied to the fourth stage blade row 33d. As a result, in the
present embodiment, it is possible to extract air having a pressure
lower than that of the air from the medium-pressure air extraction
port 16 from the low-pressure air extraction port 17 as cooling
air, and utilize this air as the cooling air for the fourth stage
blade row 33d. Thus, in the present embodiment, the compression
ratio of the cooling air for the fourth stage blade row 33d can be
decreased by the compressor 10, making it possible to suppress the
driving force of the compressor 10. Thus, in the present
embodiment, it is possible to increase the efficiency of the gas
turbine to a greater extent than in the first embodiment.
Modification of Third Embodiment
[0119] A modification of the third embodiment will be described
with reference to FIG. 11.
[0120] In the third embodiment, the end portion on the outer side
in the radial direction of the first forced vortex passage 46 is
connected to the fourth cavity 52d. The fourth cavity 52d is
connected to the second cavity 52b via the fourth disc passage 53d,
the third cavity 52c, and the third disc passage 53c, This second
cavity 52b is connected to the second stage blade row 33b via the
second stage communicating passage 55, and to the third stage blade
row 33c via the third stage communicating passage 56.
[0121] In the present modification, a bolt hole 41c of the rotor
shaft 32 through which the spindle bolt 41 is inserted is given an
elliptical cross-sectional shape, and a gap in the bolt hole 41c
through which the spindle bolt 41 is inserted serves as a passage
53e for cooling air. This passage 53e, similarly to the spindle
bolt 41, extends in the axial direction Da. The end portion on the
outer side in the radial direction of the first forced vortex
passage 46 is connected to a downstream end of this passage 53e.
Further, the second cavity 52b is connected to an upstream end of
this passage 53e. This second cavity 52b, similarly to each of the
embodiments described above, is connected to the second stage blade
row 33b via the second stage communicating passage 55, and to the
third stage blade row 33c via the third stage communicating passage
56.
[0122] Further, in the present modification, the fourth cavity 52d
is connected to the end portion on the outer side in the radial
direction of the second forced vortex passage 46b. This fourth
cavity 52d is connected to the fourth stage blade row 33d via the
fourth stage communicating passage 57.
[0123] In this way, the passage that connects the forced vortex
passage and each of the blade rows may be formed as appropriate in
accordance with a structure of the turbine rotor.
Other Modifications
[0124] While the air adjustment valves 87, 97 and the coolers 86,
86 are provided to the blade medium-pressure air extraction pipe 85
of the first, second, and third embodiments and to the blade
low-pressure air extraction pipe 95 of the third embodiment, these
are not required. Thus, the air adjustment valves 87, 97 and the
coolers 86, 96 of these air extraction pipes may be omitted, as
appropriate.
[0125] Further, in the above-described embodiments, the forced
vortex passage 46 through which cooling air for the plurality of
blade rows 33 passes is disposed on the downstream side of the
blade row 33 furthest downstream. Nevertheless, the forced vortex
passage 46 may be formed between any of the stages in the large
diameter portion 44 of the rotor shaft 32. In this case as well, at
the end on the outer side in the radial direction of the forced
vortex passage 46, this passage 46 is branched to each of the
plurality of blade rows 33, thereby making it possible to decrease
the number of forced vortex passages 46 and, as a result, suppress
elongation of the length of the rotor shaft 32 and extension of the
distance between the plurality of stages. However, the forced
vortex passage 46 is preferably formed on the downstream side of
the blade row 33 furthest downstream as in the above-described
embodiments from the viewpoint of suppressing extension of the
distance between all the stages.
[0126] Further, in each of the embodiments described above, the air
extraction ports 16, 17 are formed in the compressor casing 15, and
the air extraction line 80 is disposed outside the compressor
casing 15 and outside the turbine casing 35. Nevertheless, the air
extraction port may be formed in the compressor rotor 11, and an
axial-direction passage that extends in the axial direction inside
the compressor rotor 11 and the turbine rotor 31 may be connected
to this air extraction port, for example.
INDUSTRIAL APPLICABILITY
[0127] According to one aspect of the present invention, it is
possible to cool a blade using cooling air from a compressor and,
at the same time, suppress elongation of a rotor shaft and thus
suppress a deterioration in vibration characteristics of a
rotor.
REFERENCE SIGNS LIST
[0128] 1 Gas turbine rotor [0129] 5 Gas turbine casing [0130] 10
Compressor [0131] 11 Compressor rotor [0132] 12 Rotor shaft [0133]
13 Blade row [0134] 14 Vane row [0135] 15 Compressor easing [0136]
16 Medium-pressure air extraction port (first air extraction port)
[0137] 17 Low-pressure air extraction port (second air extraction
port) [0138] 19 Air compression flow channel [0139] 20 Combustor
[0140] 30 Turbine [0141] 31 Turbine rotor [0142] 32 Rotor shaft
[0143] 33 Blade row [0144] 33a First stage blade row [0145] 33b
Second stage blade row (first blade row) [0146] 33c Third stage
blade row (second blade row) [0147] 33d Fourth stage blade row
(third blade row) [0148] 34 Vane row [0149] 35 Turbine casing
[0150] 39 Combustion gas flow channel [0151] 43 Small diameter
portion [0152] 44 Large diameter portion [0153] 45 Axial-direction
passage (first axial-direction passage) [0154] 45b Second
axial-direction passage [0155] 46 Forced vortex passage (first
forced vortex'passage) [0156] 46b Second forced vortex passage
[0157] 47 Second stage blade array passage (first blade array
passage) [0158] 48 Third stage blade array passage (second blade
array passage) [0159] 49, 49a, 49b Fourth stage blade array passage
(third blade array passage) [0160] 52a First cavity [0161] 52b
Second cavity [0162] 52c Third cavity [0163] 52d Fourth cavity
[0164] 52e, 52f Fifth cavity [0165] 65 Large diameter portion end
cavity [0166] 66 Air receiving space [0167] 67 Pre-swirl nozzle
[0168] 70 Bearing [0169] 71 Bearing cover [0170] 80, 80b Air
extraction line [0171] 81 Medium-pressure air extraction pipe
[0172] 88 Air extraction branch pipe [0173] 91. Low-pressure air
extraction pipe
* * * * *